In a multistage compressor of a gas turbine engine, at high speed conditions the front stages usually have transonic flow conditions and carry large induced shock losses. The rear stages of the compressor have small blade heights (or span). Mechanical limitations sometimes impose large tip clearance that can result in large clearance to blade span ratio. In addition, because of rotor centrifugal effects, there may be a migration of secondary flow along blade surface from the hub to tip section of the blade, resulting in a thick tip boundary layer build up. Interaction between leading edge shocks, tip clearance vortex, blade/shroud surface boundary layer results in complex tip flow structure where low momentum flow occupies a large area near shroud. This low momentum area being accumulated downstream of these interactions may reduce rotor performance and its stall margin.